The present invention relates to an infrared (IR) suppression system, and more particularly to a rotary wing aircraft having an upwardly directed infrared suppression system which (1) masks engine exhaust IR energy, which may signal ground threats during forward flight, and (2) minimizes engine exhaust impingement on adjacent aircraft structure to reduce the overall IR signature of the rotary wing aircraft.
The exhaust ducting from a gas turbine engine is a source of high infrared energy which may be detected by heat seeking missiles and/or various forms of infrared imaging systems for targeting/tracking purposes. With respect to the former, generally speaking, a heat-seeking missile obtains directional cues from the infrared energy generated by the engine exhaust such that the amount of infrared energy given off is one of the primary determining factors of a missile's accuracy, and consequently, lethality. Regarding the latter, infrared imaging systems detect and amplify the infrared energy for detection and/or targeting.
Current IR suppression systems are utilized on many military aircraft including most rotary wing aircraft to provide IR signature reduction. Future IR threats, however, will require even greater levels of aircraft IR signature reduction.
Generally, IR suppression systems are primarily designed to: (a) reduce the infrared energy below a threshold level of a perceived threat; (b) maintain engine performance; and (c) minimize weight and packaging associated therewith. Secondary consequences may include: (i) minimizing system or configuration complexity to reduce fabrication and maintainability costs; and (ii) minimizing the external aerodynamic drag produced by such IR suppressor systems.
Current suppression systems for rotary wing aircraft are primarily designed to provide significant IR signature reduction during a hover flight profile. Generally, current suppressor systems operate by mixing the high temperature exhaust flow with cool airflow supplied by a mixing duct which communicates with an engine exhaust duct. The mixing of large amounts of ambient air with the engine exhaust may significantly reduce the overall gas temperature prior to discharging the engine exhaust overboard, thereby lowering the aircraft IR signature. To achieve significant reductions in temperature, however, a relatively significant volume of ambient air must be mixed with the high temperature exhaust flow. This requires relatively large intakes and a final exhaust stage which provides a flow area capacity for both the engine exhaust flow volume and the mixed in additional ambient airflow volume. Another disadvantage of such an IR suppressor system is limited by the packaging space restrictions. That is, the elongate mixing areas downstream of the engine need to be of a relatively significant length to provide ample mixing and flow area. Adaptation to relatively small rotary wing aircraft or retrofitting to aircraft which require maintaining current packaging constraints is therefore limited.
It is also desirable to minimize impingement of hot engine exhaust onto adjacent aircraft structure so that the generation of “hot spots” separate from the primary source associated with the nozzle/exhaust plume are avoided. Disadvantageously, the mixing operation may reduce the velocity of the exhaust flow such that the exhaust velocity may be too low to expel the exhaust far enough from the fuselage to avoid such “hot spots.” A further disadvantage is that if the exhaust gas does not have enough velocity to escape rotor downwash, the exhaust gas may be re-ingested into the engines which may reduce engine efficiency.
Accordingly, it is desirable to provide an infrared suppression system which reduces the overall IR signature of the aircraft, is compact in design, masks the IR energy emitted/radiated from the gas turbine engine for a given viewing/azimuth angle, and minimizes impingement of engine exhaust onto adjacent aircraft structure while maintaining aircraft performance characteristics.